1. Field of the Invention
The present invention relates generally to cooling air circuits of turbine rotor blades and stator vanes in gas turbine engines and, more specifically, to multiple pass serpentine cooling circuits within airfoils of the blades and vanes.
2. Discussion of the Background Art
A gas turbine engine includes a compressor that compresses air which is channeled to a combustor wherein it is mixed with fuel and ignited for generating combustion gases. The combustion gases flow downstream through one or more stages of turbines which extract energy therefrom for powering the compressor and producing additional output power for driving a fan for powering an aircraft in flight for example. A turbine stage includes a row of turbine rotor blades secured to the outer perimeter of a rotor disk with a stationary turbine nozzle having a plurality of stator vanes disposed upstream therefrom. The combustion gases flow between the stator vanes and between the turbine blades for extracting energy to rotate the rotor disk. The temperatures within gas turbines may exceed 2500 degrees Fahrenheit and cooling of turbine blades is very important in terms of blade longevity. Without cooling, turbine blades would rapidly deteriorate. Improved cooling for turbine blades is very desirable and much effort has been devoted by those skilled in the blade cooling arts to devise improved geometries for the internal cavities within turbine blades in order to enhance cooling. Since the combustion gases are hot, the turbine vanes and blades are typically cooled with a portion of compressor air bled from the compressor for this purpose. Diverting any portion of the compressor air from use in the combustor necessarily decreases the overall efficiency of the engine. Accordingly, it is desired to cool the vanes and blades with as little compressor bleed air as possible.
Typical turbine vanes and blades include an airfoil over which the combustion gases flow. The airfoil typically includes one or more serpentine cooling passages therein through which the compressor bleed air is channeled for cooling the airfoil. The airfoil may include various turbulators therein for enhancing cooling effectiveness and the cooling air is discharged from the passages through various film cooling holes disposed around the outer surface of the airfoil. In pursuit of higher cooling effectiveness, modern blades have led to multi-pass cooling circuits with many cavities such as 5 passes and 5 cavities. A drawback to having more cavities is that more ribs are required which result in more weight. At some combinations of scale and rotational speed, the heavier blade designs result in heavy rotor disks which are difficult to design for long life. Also, the use of a multi-pass serpentine requires extra coolant supply pressure. If the required coolant supply pressure can be reduced, then cooler air from an earlier compressor stage can be used. This benefits cycle performance since less work is put into the coolant.
Known turbine airfoil cooling techniques include the use of internal cavities forming a serpentine cooling circuit. Particularly, serpentine passages, leading edge impingement bridges, film holes, pin fins, and trailing edge holes or pressure side bleed slots are utilized for blade cooling. It would be desirable to provide improved blade cooling. In providing even better blade cooling, it also would be desirable to avoid significantly increasing the blade fabrication costs.
A gas turbine engine hollow airfoil with an airfoil outer wall having transversely spaced apart pressure and suction side walls joined together at chordally spaced apart leading and trailing edges of the airfoil and extending radially from a base to a tip. Inside the airfoil is a plurality of radially extending internal ribs extending width wise between the pressure and suction side walls and a single internal aft flowing three pass serpentine cooling circuit having radially extending first, second, and third serpentine channels between, in axially aft succession, first, second, third, and fourth ribs of the radially extending internal ribs. The serpentine cooling circuit has terminal end that is positioned aft of the entrance so as to have a chordal flow direction aftward from the leading edge to the trailing edge within the serpentine circuit. A straight through leading edge feed channel extends radially through the airfoil and bounded in part by the leading edge. The exemplary embodiment includes a trailing edge cooling plenum and a straight through single pass trailing edge feed channel for providing cooling air to the trailing edge cooling plenum. The airfoil is on a turbine blade in the exemplary embodiment of the invention illustrated herein.
The present invention provides advantages that include a good cooling of a hollow gas turbine airfoil using less cooling air than would otherwise be necessary while still providing sufficient cooling for the airfoil tip and acceptable airfoil tip metal temperatures. The cooling air in the mid-circuit can be tailored for the pressure side wall heat load, thus, allowing cooler temperatures at the tip of the last up-pass and better tip cooling. The downstreamwise serpentine circuit design of the present invention provides the coldest cooling air in the hottest areas of the blade. The cooling air temperatures are colder than the cooling air temperatures in the same channels and chambers in conventional upstreamwise serpentine circuit designs. The downstreamwise serpentine circuit will have a colder average spanwise rib wall temperature than that of an upstreamwise serpentine circuit and, therefore, have an overall better cooling air temperature distribution in the chordwise direction and a better bulk temperature of the airfoil for better cooling of the entire airfoil.
The leading edge is cooled by colder fresher cooling air than in those in the prior art reducing or eliminating the amount of film cooling required in this region and the straight through single pass channel used to cool the leading edge reduces pressure losses associated with impingement leading edge cooling chambers. In addition, the film cooling holes closer to the trailing edge can have shallower flow angles from surface than those closer to the leading edge resulting in a better film cooling effectiveness. The external gas flow velocity closer to the trailing edge accelerates to a higher speed than at portions along the airfoil side walls closer to the leading edge. Therefore, the airfoil cooling can be better tailored for conductive and convective cooling of portions of the sides of the outer wall closer to the leading edge and film cooling holes may be used for portions of the sides closer to the trailing edge where they will have smaller and, therefore, better blowing ratios and result in a better film cooling effectiveness and overall cooling efficiency.
Other advantages include increased coolant side heat transfer coefficient and improved metering capability for external film flow. The improved cooling also provides for cooler air to be discharged through the tip cooling holes providing improved cooling for the squealer tip.
Generally, design requirements for airfoils at the lower spans are driven by concerns for rupture at high stress levels at reduced metal temperature and at the upper regions by concerns over elevated surface temperature to avoid oxidation and fatigue crack initiation. The downstream flowing serpentine flow channels of the present invention addresses these needs along with the ability to better optimize internal airfoil cooling flow and blade life.
The cooling circuit configuration of the present invention allows the use of a lower coolant supply pressure. The three pass serpentine is also less vulnerable to variations in pressure drops from cast features than the circuits having more channels and passes. Dedicated circuits or channels for leading edge and trailing edge cooling provide better internal cooling at the edges where the external heat load is highest. The straight through separate leading edge channel helps the blade be tolerant of holes from foreign object damage. The impingement cavity at the trailing edge allows good support of the aftmost cavity core during the casting process. This cavity tends to be thin where the airfoil shape is tapering toward the trailing edge.
The present invention is capable of providing good cooling of the hollow gas turbine airfoil using less cooling air than would otherwise be necessary while providing an even distribution of temperatures for reduced thermal stresses. The aft flowing serpentine cooling circuit provides convective cooling in the region where the leading edge circuit can provide film coverage. The last pass of the serpentine then feeds film holes to cover the region where the trailing edge circuit provides convection. By allowing the film to exit at the aftmost part of the serpentine, the invention improves the film cooling benefit at the trailing edge where convection cooling is difficult.